Radial flow, gas turbine engines have long been known and employed in a variety of applications. Quite obviously, they are desirable for those applications where axial compactness is required. They have also seen extensive use in applications where their relative simplicity is highly desirable as, for example, in auxiliary power units for aircraft. Indeed, they have even been employed in thrust jet applications for small aerodynamic vehicles as, for example, drones and cruise missiles.
Many of today's radial flow, gas turbine engines employ so-called "monorotors" wherein the monorotor consists of a single rotating mass defining both the compressor section and the turbine wheel section of the engine. Because the rotor must stand up to the high temperatures encountered in the hot section of the engine, i.e., the temperatures to which the turbine wheel section of the monorotor is exposed, it is made of a single alloy. Typically, the alloy employed is a cast nickel based superalloy which will accommodate the elevated temperature that is experienced in the hot section. Alloys such as nickel based superalloys or the like, while providing excellent resistance to the high temperatures encountered, have substantial density and thus result in a monorotor having a relatively high mass. The relatively high mass of the monorotor results in relatively large inertia which in turn makes the engine more difficult to start than would be the case if the inertia of the rotor was less. Similarly, the high mass of such monorotors results in relatively high loading of the engine bearings. Such loading can, in turn, require larger bearings than would otherwise be necessary and/or result in a shorter useful life of the gas turbine engine and/or necessitate the use of improved lubrication systems.
All of the foregoing contributes to the cost of the gas turbine engine, as well as affects its reliability in terms of ease of starting, longevity, etc.
The present invention is directed to overcoming one or more of the above problems.